1. Field of the Invention
The present disclosure relates to gas turbine engines, and more particularly to vibration dampers for gas turbine engine compressor and turbine disk blade assemblies.
2. Description of Related Art
Gas turbine engines typically include one or more compressor and turbine disk assemblies. The disk assemblies typically include a disk portion with disk slots defined about the circumference of the disk with blades seated in the slots. Some gas turbine engines include blade dampers positioned between the roots of adjacent blades between the undersides of adjacent blade platforms and the disk portion. Such blade dampers typically dampen the first vibratory mode of the airfoil during engine operation.
The dampening effect of conventional blade dampers is generally a function of the orientation of the blade damper in relation to the adjacent blades. Obtaining a desired or predetermined damping effect generally requires that the damper be installed in one or a limited number of orientations in order to provide a desired damping effect to the blades. In some engine designs the blade damper can be installed in an orientation where it does not provide the desired damping effect, potentially requiring removal and reinstallation of the blade damper such that it is installed in its intended orientation. Disk assemblies with relatively small blade dampers, such as high-pressure turbine disks, can be particularly susceptible to assembly errors due to blade damper orientation.
Such conventional turbine dampers have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved turbine dampers that simplify engine assembly. The present disclosure provides solutions to this need.